U.S. patent number 5,984,236 [Application Number 08/577,444] was granted by the patent office on 1999-11-16 for momentum unloading using gimbaled thrusters.
Invention is credited to David K. Abernethy, Bernard M. Anzel, Keith F. Keitel, Richard A. Noyola, John F. Yocum, Jr..
United States Patent |
5,984,236 |
Keitel , et al. |
November 16, 1999 |
Momentum unloading using gimbaled thrusters
Abstract
A method of simultaneously controlling East/West and North/South
positioning and unloading momentum of a spacecraft while orbiting
an object. The spacecraft has a thruster array and a momentum
accumulator. The method entails moving said spacecraft towards a
node of the orbit. At a predetermined position on the orbit,
separate from the node, a thruster of the thruster array is fired
so as to control the orbital position of the spacecraft. While the
thruster is being fired, momentum is dumped from the momentum
accumulator at the predetermined position so that any loss in
control in the attitude of the spacecraft is reduced.
Inventors: |
Keitel; Keith F. (Los Angeles,
CA), Noyola; Richard A. (Torrance, CA), Yocum, Jr.; John
F. (Rancho Palos Verdes, CA), Abernethy; David K. (Los
Angeles, CA), Anzel; Bernard M. (El Segundo, CA) |
Family
ID: |
24308770 |
Appl.
No.: |
08/577,444 |
Filed: |
December 22, 1995 |
Current U.S.
Class: |
244/164; 244/169;
244/165; 244/171.2 |
Current CPC
Class: |
B64G
1/283 (20130101); B64G 1/26 (20130101); B64G
1/24 (20130101); B64G 1/285 (20130101); B64G
1/28 (20130101) |
Current International
Class: |
B64G
1/24 (20060101); B64G 1/26 (20060101); B64G
1/28 (20060101); B64G 001/26 () |
Field of
Search: |
;244/168,164,165,166,171,158R ;364/459,429.023 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Jordan; Charles T.
Assistant Examiner: Dinh; Tran
Attorney, Agent or Firm: Gudmestad; Terje Grunebach;
Georgann Sales; Michael W.
Claims
We claim:
1. A method of simultaneously controlling East/West and North/South
positioning and unloading momentum of a spacecraft while orbiting
the Earth, said spacecraft having a thruster array, an on-board
processor capable of receiving stationkeeping commands from a
ground station located on the Earth that correspond to desired
spacecraft velocity changes and capable of calculating thruster
firing parameters necessary to meet both stationkeeping and
momentum dumping requirements of said spacecraft, and a momentum
accumulator, said method comprising the steps of:
receiving desired spacecraft velocity changes from said ground
station;
moving said spacecraft towards a node of said orbit;
calculating, solely using said on-board processor, thruster firing
parameters necessary to meet said stationkeeping requirements and
momentum dumping requirements;
firing a thruster of said thruster array at a predetermined
position on said orbit, separate from said node, so as to control
the orbital position of said spacecraft, wherein said predetermined
position is calculated by said on-board processor; and
simultaneously dumping momentum from said momentum accumulator at
said predetermined position and while said thruster is being fired
so that any loss in control in the attitude of said spacecraft is
reduced.
2. The method of claim 1, wherein said step of dumping momentum
comprises gimballing a thruster of said thruster array.
3. The method of claim 2, wherein said fired thruster and said
gimballed thruster are one and the same.
4. The method of claim 1, further comprising the step of
continuously monitoring momentum stored by said momentum
accumulator.
5. The method of claim 1, wherein said spacecraft comprises a first
momentum dumper and a second momentum dumper, said method
comprising the steps of:
continuously monitoring the amount of momentum stored by said
momentum accumulator; and
wherein said step of dumping momentum comprises dumping momentum
from one or both of said first and second momentum dumpers in
response to the amount of momentum stored by said momentum
accumulator.
6. The method of claim 1, wherein the location of said
predetermined position is determined based upon a predetermined
amount of momentum which needs to be dumped from said spacecraft
during said momentum dumping step.
7. The method of claim 6, wherein said predetermined position is
located approximately 30.degree. from said node.
8. A method of simultaneously controlling East/West and North/South
positioning and unloading momentum of a spacecraft while orbiting
the Earth, said spacecraft having first, second, third and fourth
thrusters, an on-board processor capable of receiving
stationkeeping commands from a ground station located on the Earth
that correspond to desired spacecraft velocity changes and capable
of calculating thruster firing parameters necessary to meet both
stationkeeping and momentum dumping requirements of said
spacecraft, and a momentum accumulator, said method comprising the
steps of:
moving said spacecraft towards a node of said orbit;
calculating, solely using said on-board processor, thruster firing
parameters necessary to meet said stationkeeping requirements and
momentum dumping requirements;
firing said first thruster at a predetermined position on said
orbit, separate from said node, so as to control the orbital
position of said spacecraft, wherein said predetermined position is
calculated by said on-board processor; and
simultaneously dumping momentum from said momentum accumulator at
said predetermined position and while said first thruster is being
fired so that any loss in control in the attitude of said
spacecraft is reduced.
9. The method of claim 8, wherein the location of said
predetermined position is determined based upon a predetermined
amount of momentum which needs to be dumped from said spacecraft
during said momentum dumping step.
10. The method of claim 9, wherein said first, second, third and
fourth thrusters are arranged at a northwest corner, a northwest
corner, a southwest corner and a southeast corner, respectively, of
a rectangular array.
11. The method of claim 10, further comprising the steps of:
moving said spacecraft past said node of said orbit;
firing said second thruster at a second predetermined position on
said orbit, separate from said node, so as to control the orbital
velocity and attitude of said space-craft; and
dumping momentum from said momentum accumulator at said second
predetermined position and while said second thruster is being
fired so that any loss in control in the attitude of said
spacecraft is reduced.
12. The method of claim 11, wherein said second predetermined
position is located up to 30.degree. from said node.
13. The method of claim 11, further comprising the steps of:
moving said spacecraft towards a second node of said orbit;
firing said third thruster at a third predetermined position on
said orbit, separate from said second node, so as to control the
orbital velocity and attitude of said space-craft; and
dumping momentum from said momentum accumulator at said third
predetermined position and while said third thruster is being fired
so that any loss in control in the attitude of said spacecraft is
reduced.
14. The method of claim 13, wherein said third predetermined
position is located up to 30.degree. from said second node.
15. The method of claim 13, further comprising the steps of:
moving said spacecraft past said second node of said orbit;
firing said fourth thruster at a fourth predetermined position on
said orbit, separate from said second node, so as to control the
orbital position of said spacecraft; and
dumping momentum from said momentum accumulator at said fourth
predetermined position and while said fourth thruster is being
fired so that any loss in control in the attitude of said
spacecraft is reduced.
16. The method of claim 15, wherein said fourth predetermined
position is located up to 30.degree. from said second node.
17. The method of claim 8, wherein said predetermined position is
located up to 30.degree. from said node.
18. A method of simultaneously controlling East/West and
North/South positioning and unloading momentum from a momentum
accumulator on a spacecraft, said spacecraft including an on-board
processor capable of receiving stationkeeping commands from a
ground station located on the Earth that correspond to desired
spacecraft velocity changes and capable of calculating thruster
firing parameters necessary to meet both stationkeeping and
momentum dumping requirements of said spacecraft, a first momentum
dumper and a second momentum dumper, said method comprising the
steps of:
monitoring the amount of momentum stored by said momentum
accumulator;
calculating thruster firing parameters necessary to meet both
stationkeeping and momentum dumping requirements of said spacecraft
solely by utilizing said on-board processor; and
activating one or both of said first and second momentum dumpers to
simultaneously dump momentum from said momentum accumulator in
response to the momentum dumping requirement calculated by said
on-board processor.
19. The method of claim 18, wherein said first momentum dumper
comprises a thruster.
20. The method of claim 18, wherein said second momentum dumper
comprises a gimbaled solar panel.
21. The method of claim 18, wherein said second momentum dumper
comprises a magnetic torquer.
Description
BACKGROUND OF THE INVENTION
Currently, when a spacecraft, such as a satellite, is moving in an
orbit about a planetary body, such as the Earth, the spacecraft
encounters disturbances, such as solar wind. These disturbances,
left unchecked, will create a series of impulses of momentum which
will saturate the momentum accumulator which could cause a loss in
attitude. The amount of deviation is directly related to the
specific configuration of the spacecraft--the more symmetric a
spacecraft is about its center of mass the amount of deviation is
minimized. For example, the most severe configuration of satellite
with a single reflector can result in the satellite encountering
daily momentum accumulations in roll, pitch and yaw by amounts of
5, 10 and 5 Nms, respectively. In a more typical case of a
satellite having two symmetrically mounted reflectors, momentum
accumulation in roll, pitch and yaw can each amount to 5 Nms.
To counteract such accumulations of momentum, spacecraft employ
momentum accumulators which store the momentum encountered by the
spacecraft so that the effect of the momentum is minimized or
reduced. Examples of well known momentum accumulators is a pyramid
of reaction wheels or gimbaled momentum wheels.
These momentum accumulators, however, are unable to accumulate
momentum without end. Eventually, the stored momentum needs to be
dumped or unloaded during the orbit of the spacecraft. However, the
manner in which a spacecraft performs thruster operations to remain
in a desired orbit, known as stationkeeping, does have an effect on
the momentum dumping capability of the spacecraft. Furthermore, in
such stationkeeping, the dumping capability is maximized when the
total daily burn time is maximum (maximum inclination delta-v) and
when the burn time is distributed most symmetrically throughout the
day (minimum eccentricity and longitudinal acceleration delta-v).
The worst-case conditions for dumping capability are therefore the
minimum-eccentricity control strategy with minimum north-south
disturbance in the 17 year life cycle of a satellite and with
maximum longitudinal acceleration.
Several stationkeeping methodologies are possible. For example,
prior satellites have unloaded momentum simultaneous with
North/South stationkeeping only by using thrusters mounted on the
North face. Simultaneous control of East/West positioning is not
contemplated with this method.
Another possible method of stationkeeping is described in U.S. Pat.
No. 5,443,231 to Anzel. That application describes a method of
East/West and North/South stationkeeping which uses four gimballed
ion thrusters in the same configuration as shown in FIG. 2 of the
present application.
In a third example, gimballed ion thrusters are mounted on the
North face and the South face of a satellite for momentum unloading
during North/South stationkeeping, such as described in U.S. Pat.
No. 5,349,532 to Tilley et al. Again, East/West stationkeeping is
not performed.
Furthermore, gimballed ion thrusters are used for North-South
stationkeeping on the EUROSTAR Spacecraft. This spacecraft appears
not to disclose momentum unloading nor the use of the system for
East-West stationkeeping.
While the above-mentioned control systems are generally adequate
for their intended purpose, there is room for improvement. For
example, the fuel efficiency of the above-mentioned control systems
is adversely affected because North/South and East/West
stationkeeping and momentum dumping are not performed
simultaneously. Furthermore, unloading of momentum is not done in
an efficient manner by prioritizing which systems are activated to
dump momentum.
SUMMARY OF THE INVENTION
The present invention provides a control system for controlling the
orbital position of a spacecraft. The control system of the present
invention is capable of controlling the orbital position of the
spacecraft while simultaneously dumping momentum so as to increase
the fuel efficiency of the spacecraft. In particular, the present
invention concerns a method of simultaneously controlling East/West
and North/South positioning and unloading momentum of a spacecraft
while orbiting an object. The spacecraft has a thruster array and a
momentum accumulator. The method entails moving said spacecraft
towards a node of the orbit. At a predetermined position on the
orbit, separate from the node, a thruster of the thruster array is
fired so as to control the orbital position of the spacecraft.
While the thruster is being fired, momentum is dumped from the
momentum accumulator at the predetermined position so that any loss
in control in the orbital position is reduced.
Another aspect of the present invention regards a method of
simultaneously controlling East/West and North/South positioning
and unloading momentum of a spacecraft while orbiting an object.
The spacecraft has four thrusters and a momentum accumulator. The
method entails moving the spacecraft towards a node of the orbit.
One of the thrusters is fired at a predetermined position on the
orbit, separate from the node, so as to control the orbital
position of the spacecraft. While the thruster is being fired at
the predetermined position, momentum is dumped from the momentum
accumulator so that any loss in control in the orbital position is
reduced.
A third aspect of the present invention is a spacecraft control
system for simultaneously controlling East/West and North/South
positioning and unloading momentum of a spacecraft while orbiting
an object. The spacecraft control system includes
a spacecraft having a thruster array and a momentum accumulator.
The system further includes a sensor for generating a signal
representative when the spacecraft has arrived at a predetermined
position on the orbit which is separate from a node of the
orbit.
A momentum controller is provided which receives the signal and
sends a signal to a thruster of the thruster array so that the
thruster fires at the predetermined position in response to receipt
of the signal so as to control the orbital position of the
spacecraft. The momentum controller also sends a signal to the
momentum accumulator to dump momentum at the predetermined position
and while the thruster is being fired so that any loss in control
in the orbital position is reduced.
The above-described control system of the present invention
improves fuel efficiency on the spacecraft which allows a
spacecraft to perform stationkeeping solely via ion propulsion
thrusters and, thus, reduces the need for bipropellant fuel or
thrusters. The disclosed control system efficiently unloads
momentum by prioritizing which systems are activated to dump
momentum.
The foregoing features and advantages of the present invention will
be further understood upon consideration of the following detailed
description of the invention taken in conjunction with the
accompanying drawings, in which:
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 shows a perspective view of a spacecraft capable of
utilizing a stationkeeping and momentum dumping system and method
of the present invention;
FIG. 2 shows a thruster configuration to be used with the
spacecraft of FIG. 1;
FIG. 3 shows a gimbaled thruster to be used with the thruster array
of FIG. 2;
FIG. 4 shows an orbital burn sequence for the southwest and
southeast burns for the spacecraft of FIG. 1;
FIG. 5 schematically shows a stationkeeping and momentum dumping
control system for the spacecraft of FIG. 1;
FIG. 6 shows the gimbal angle required to dump 5 Nms or roll/yaw
momentum as a function of pitch momentum dumped daily;
FIG. 7 shows a typical profile of burn durations throughout the
year;
FIG. 8 shows the roll/yaw dumping capability at each of the two
nodes and the effect of burn time distribution on dumping
capability;
FIG. 9 shows a comparison of two dump strategies of the present
invention; and
FIG. 10 shows a comparison of the two dump strategies where the
transient momentum is more apparent.
DESCRIPTION OF THE PREFERRED EMBODIMENT
A spacecraft control system for controlling the stationkeeping and
momentum dumping or unloading of a spacecraft or satellite 20
according to the present invention is substantially shown in FIGS.
1-10. Satellite 20 is a three axis-stabilized spacecraft having a
spacecraft body 22 which includes a lower bus module 24 and an
upper payload module 26. Attached to the aft end of the lower bus
module 24 are a plurality of engines which will be discussed in
detail later. Lower bus module 24 contains fuel tanks (not shown)
and various power and control modules which operate the engines and
power the payload module 26. Bus module 24 further includes a pair
of solar panels 28 which convert sunlight into electricity which is
sent to batteries (not shown) located on the bus module 24. Bus
module 24 also has a pair of antennae 30, which receive signals
from a ground station on Earth which are used to control the
satellite. Antennae 30 also send signals to the ground station.
Payload module 26 is attached to the bus module 24 and contains a
variety of electronic equipment which may contain a number of
sensors 32. The electronic equipment processes information gathered
by sensors 32 and sends the processed information back to the
ground station via antennae 30. Payload module 26 further includes
heat radiators 34 which emit heat generated by the satellite
20.
As shown in FIGS. 1 and 2, spacecraft or satellite 20 has four
thrusters 36, 38, 40, 42 mounted on the anti-nadir side of bus
module 24. The orbit inclination, eccentricity, and rate mean
motion of spacecraft 20 can be controlled by firing all of the
thrusters 36, 38, 40 and 42 per day, where only one thruster is
fired at a time at predetermined times and duration. All four of
the thrusters are aligned such that their nominal thrust vectors
are directed through the spacecraft center of mass as described in
U.S. Pat. No. 5,443,231 to Anzel, whose contents are incorporated
herein by reference. Each of the north and south thruster pairs are
mounted on a thruster platform which includes a single two-axis
gimbal mechanism 44 (see FIG. 3) such that the thrust vectors make
an angle .theta. (typically 45.degree.-55.degree.) with the
north-south axis being in a northerly and southerly direction,
respectively. The thrusters are also separated in the east-west
direction by an angle a (typically 10.degree.-13.degree.). Each
thruster can produce a component of force or torque about all three
axes. The thrusters can be gimballed to produce torques necessary
to unload momentum from any storage devices, while simultaneously
controlling East/West and North/South positioning of the
spacecraft. Besides gimbaled thrusters 36, 38, 40 and 42, momentum
stored in momentum accumulators 45, such as reaction or momentum
wheels, can be dumped or unloaded by moving solar panels 28 or
employing well known magnetic torquers 46 (schematically shown in
FIG. 1). It should be noted that the momentum accumulators are the
primary pointing control actuators.
As shown in FIG. 3, each thruster is attached to a gimbaled
mechanism 44 and a pair of actuators 47, 48. Each actuator 47, 48
can employ stepper motors mounted at right angles from each other
so that the actuators 47, 48 can change their lengths in response
to a control signal so as to control the angled position of the
gimbaled thrusters 36, 38, 40, 42 in two orthogonal directions,
.theta., .PHI.. Each of the four thrusters 36, 38, 40 and 42
preferably is an electronic thruster, such as a xenon ion
propulsion thruster.
FIG. 5 schematically shows the structure of a momentum management
system 50 which may utilize several actuators to counteract
disturbance torques (such as from solar flux on the spacecraft) and
to unload momentum from the momentum accumulators or storage
devices 45. The momentum management system 50 ideally will be able
to formulate and execute a strategy for using these burns, as
defined from stationkeeping requirements, to dump the necessary
disturbance momentum from the spacecraft with minimal impact to the
stationkeeping capabilities and no additional fuel cost. Momentum
management system 50 preferably comprises one or more
microprocessors to perform the calculations and tasks outlined
below. Furthermore, the entire momentum management system 50 can be
located on board the spacecraft 20.
A first step in managing momentum on spacecraft 20 is for a ground
station on the Earth to send commands 52 to satellite 20 which
correspond to daily velocity changes, .DELTA. V, to be performed
over a period of time, such as a two week period, for
stationkeeping. The .DELTA. V changes are calculated at the ground
station or on the spacecraft in a manner specified in U.S. Pat. No.
5,443,231. The calculations are based, in part, on feedback signals
56 sent from the satellite 20 to the ground station which represent
the orbital dynamics of the satellite 20 in the previous two
weeks.
Preferably, it would be desired that the satellite 20 would have
sufficient processor ability to calculate both the stationkeeping
and momentum dumping requirements of the satellite 20 without input
from the ground station. This increases the robustness of the
stationkeeping control system in the presence of changing
disturbances and torques.
The .DELTA. V changes sent to the satellite 20 are received by
antennae 30 and relayed to a .DELTA. V buffer 58 which stores
fourteen days of .DELTA. V changes. The values stored correspond to
the variables .DELTA. Vin, .DELTA.Ver, .DELTA. Vet and .DELTA.Vdt
which are described in U.S. Pat. No. 5,443,231, whose contents are
incorporated herein by reference.
During the continuous operation of satellite 20, a momentum manager
60 monitors each of the on-board momentum accumulators and computes
momentum unloading commands .DELTA. H.sub.1, .DELTA. H.sub.2, and
.DELTA. H.sub.3 which represent the total momentum to be dumped in
a 24 hour period along the X, Y and Z directions, respectively.
One method of calculating .DELTA. H.sub.1, .DELTA. H.sub.2, and
.DELTA. H.sub.3 is given below. This method is based on minimizing
the thruster platform's gimbal extremes. This minimization is
achieved at the cost of a higher transient momentum storage between
burns, thereby increasing the necessary momentum margin which must
be used in sizing the momentum wheels.
Four burns, one burn per thruster, should be made during the travel
of the satellite 20 along its orbit, if no thrusters have failed.
The position of the burns is specified by momentum control 62. As
shown in FIG. 4, as the satellite 20 nears the descending node of
the orbit it will fire its southwest thruster 38 for a duration of
time D.sub.1 and at an angular position .epsilon..sub.w. Similarly,
there will be three other burns. One burn will involve firing the
southeast thruster 40 for a duration D.sub.2 on the other side of
the descending node at .epsilon..sub.e. After the second burn, the
satellite 20 moves towards the ascending node where a burn of the
northwest thruster 36 is performed at .epsilon..sub.w for duration
D.sub.3. After the satellite passes the ascending node, the final
burn for the 24 hour period is performed at .epsilon..sub.e where a
northeast thruster 42 is fired for a duration D.sub.4. The values
for .epsilon..sub.e, .epsilon..sub.w, D.sub.1, D.sub.2, D.sub.3,
D.sub.4 are discussed in U.S. Pat. No. 5,443,231, whose contents
are incorporated herein by reference. FIG. 6 shows the gimbal
angles as a function of pitch momentum dumped daily. These results
assume that the spacecraft is not offset in pitch.
As shown in FIG. 7, the two quantities (D.sub.3 +D.sub.1) and
(D.sub.4 +D.sub.2) are constant throughout the year. The resulting
geometry allows for a relatively constant momentum unloading
capability throughout the year. There is an effect for the
techniques described below for when a thruster fails or when the
roll transient storage is minimized. Note that the daily total burn
time (nominally 30 minutes) does not vary significantly during the
year, but the duration of each thruster burn may vary from 3
minutes to 20 minutes through the year.
The total momentum components .DELTA. H.sub.1, .DELTA. H.sub.2, and
.DELTA. H.sub.3 are calculated as the sum of the current momentum
accumulator momentum, the expected momentum growth from secular
disturbance torques, and a momentum target command which will
center the momentum variations over the next 24-hour period. These
components are always computed in the frame coincident with the
spacecraft body axes at the node which is located an angular amount
.PHI. away from the time when the calculation is made. Accordingly,
.DELTA. H.sub.1, .DELTA. H.sub.2, and .DELTA. H.sub.3 are
calculated below: ##EQU1## where, the H.sup.now components
represent the momentum stored in the accumulators at position
.PHI., the L.sup.d terms are estimated torque components for a 24
hour period and are received from estimator 48. .DELTA.
H.sup.Target represents a factor which removes any bias stored in
the spacecraft 20 during the day and reduces momentum storage
requirements. The calculation of .DELTA. H is preferably done twice
each day before each node. In addition, ##EQU2## where: s.sub.w
=sin(.alpha.+.epsilon..sub.w), s.sub.e
=sin(.alpha.+.epsilon..sub.e, c.sub.w
=cos(.alpha.+.epsilon..sub.w), c.sub.e
=cos(.alpha.+.epsilon..sub.e). .alpha. and .theta. are the nominal
thruster angles which are normally 13.degree. and 45.degree.,
respectively. Furthermore, the distribution factors: k.sub.w
=D.sub.w.sup.now /(D.sub.w.sup.now +D.sub.w.sup.next), k.sub.e
=D.sub.e.sup.now /(D.sub.e.sup.now +D.sub.e.sup.next), where
D.sub.w and D.sub.e denotes D.sub.1 and D.sub.2 respectively if at
the descending node (southwest and southeast burn durations) or
D.sub.3 and D.sub.4 respectively if at the ascending node
(northwest and northeast burn durations). The superscripts now and
next refer to the burn durations at the present burn location and
the next burn location, respectively.
One way to dump the accumulated momentum is to use gimballed
thrusters 36, 38, 40 and 42. Preferably, the gimballed thrusters
each use 3-jackscrew mechanism that provides a 3-for-2 redundancy
should one of the jackscrews fail. Each thruster can be gimballed
in two axes about its thrust vector (up to .+-.14.degree.) so that
each thruster burn can produce dumping torques orthogonal to the
thrust vector. This dumping plane can nominally be described with
respect to the spacecraft frame {x,y,z} as the plane orthogonal to
the vector (y-z) for the south thrusters 38, 40 and a plane
orthogonal to the vector (y+z) for the north thrusters 36, 42.
Inertially, however, because the ascending (North) and descending
(South) burns are separated by 12 hours, these planes are nearly
coincident, and the dumping capability is severely limited along
the inertial thrust vector.
Roll torques (with minimum yaw coupling) are produced by gimballing
the thrusters in the North-South direction by an angular amount
.rho.. Gimballing in the East-West direction by an angular amount
.gamma., produces a combination of pitch and yaw torques (with
minimum roll coupling). The capability for dumping momentum along
the inertial thrust vector relies on the East-West slew angle
(typically 10.degree.-13.degree.) of the thrusters 36, 38, 40, 42.
Such an approach requires a roll torque during the northwest burn
which is opposite in polarity from the roll torque during the
northeast burn, such that the sum of the torques sum produce a yaw
torque in the inertial frame. This capability is reduced, however,
by the yaw torques produced by the pitch dumping. This yaw dump
capability can be increased significantly by moving the burns away
from the node by perhaps up to 30.degree. (which effectively
increases the coupling of roll torques into yaw).
Furthermore, assuming that the roll/yaw dumping is distributed
between the nodes based on roll/yaw dumping capability to minimize
the gimbal angle, and pitch dumping is distributed between the
nodes to allow for removal of entire yaw by-product momentum at the
same node, then .DELTA. H can be expressed as follows: ##EQU3##
this can be reduced to the following matrix form: ##EQU4## Note
that this distribution equation assumes that the roll/yaw byproduct
from pitch dumping is purely along the z-axis. This momentum
component may diverge from the z-axis by up to an angular amount
(.alpha.+.epsilon.) depending on burn distribution and
.delta..sub.yz as described below. Also, note that FIG. 8 shows the
roll/yaw dumping capability at each of the two nodes and the effect
of burn time distribution on dumping capability.
The calculated values for .DELTA. H and the stored .DELTA. V values
are fed to a momentum control 62 where the burn parameters
.epsilon..sub.e, .epsilon..sub.w, D.sub.1, D.sub.2, D.sub.3,
D.sub.4 and the torque commands L at each burn are calculated.
Since the calculation of the torque commands and burn duration
times depend on the value of the other, an iteration is performed
to calculate both quantities. The final calculated values for the
burn durations and the torque commands are those which give stable
solutions during the iteration process. The torque commands are for
torques about 2-axes in the satellite frame which are computed such
that gimbal angles .rho.,.gamma. are minimized. The burn commands
or durations are sent to a thruster sequencer 64 which uses the
burn commands to send commands to the thrusters 36, 38, 40, 42
where the signals control the fuel valves and thruster power for
each thruster. The torque commands are sent to a momentum dumping
controller 66. In the case of dumping momentum via the gimballed
thrusters, controller 66 uses the torque commands in combination
with a torque feedback signal generated by an estimator 68. The
torque feed back signal is representative of an estimate of the
torque error in the thrusters. Estimator 68 estimates the torque
error based on the attitude of the spacecraft 20 as determined by
an internal reference unit (IRU) 70 and a spacecraft dynamics
processor 72. The IRU 70 typically consists of a gyro which acts as
a rate sensor for estimating the body and angular rates of the
spacecraft 20.
The spacecraft dynamics processor contains data regarding the
physical characteristics of the spacecraft 20 and calculates the
attitude and orbital dynamics of the spacecraft 20 based on
receiving a torque signal 74 representative of the torques 76 and
78 generated by the thrusters and the gimbaled thruster platform,
respectively, and extraneous torques 80 exerted on the spacecraft
20, which include torques caused from solar, radio-frequency,
thruster platform stepping, thrust variation and burn timing
factors. Besides the IRU 70, the spacecraft dynamics processor 72
sends a spacecraft dynamics signal 56 to the ground station to be
processed by the Kalman filter 54, as mentioned previously. Note
that the thruster torque signal 76 is fed back to thruster
sequencer 64 so that the telemetry of the thrusters can be
monitored and used to adjust the burn durations for arcing.
Momentum control 62 determines the torque commands from the
relationship below: ##EQU5## This can be rewritten as follows,
because .delta..sub.yz is determined a priori in a well known
manner: ##EQU6## where: pitch differential torque, .delta..sub.yz
is nominally zero and can be used to reduce the .rho. gimbal angle
at the cost of increased gimbal angle .gamma. (see below), and
nominal xenon thruster platform cant angle, .theta.=45.degree. for
D.sub.3 & D.sub.4 burns, and .theta.=-45.degree. for D.sub.1
& D.sub.2 burns.
All the previous calculations for momentum and torques are
performed at a position .PHI. (see FIG. 4) located prior to
reaching the burn location. However, just prior to the burn
location a mathematical transformation of the torques is performed.
Since the above-mentioned 2-axis torque commands have been
calculated in the thruster frame, they should be resolved once for
each burn into 3-axis torque commands (in the spacecraft frame).
This transformation is given below: ##EQU7## where burn number, i,
is equal to 1 or 2. When i=1 that represents that a torque is for
either a northwest burn or a southwest burn. Similarly, i=2
represents the torque for either a northeast or a southeast
burn.
Also, just prior to the burn location, the thruster platform is
stepped into position and a closed-loop gimbal command is performed
throughout each burn to null torque error, using a 2.times.3
pseudo-inverse transformation matrix. A good approximation is shown
below in which the quantities are in deg/Nm and would be calibrated
once in each orbit: ##EQU8##
While a burn is performed, there is a gimbal command which limits
the calculated values for the gimbal angles to lie between
+/-.gamma..sub.max, .rho..sub.max, where .gamma..sub.max,
.rho..sub.max represent the maximum gimbal angles which are
structurally possible for the gimbaling mechanisms on the thruster
platform.
Momentum dumping control 66 also prioritizes or controls which
momentum accumulators are dumped or momentum dumpers are activated,
when the momentum accumulators are dumped, when the momentum
dumpers are activated and how much momentum is dumped by each
momentum accumulator or momentum dumper. In order to accomplish
this level of control, the momentum dumping control 66 will
prioritize and weight the commands sent to each momentum
accumulator. Furthermore, the momentum accumulators may employ
either bang-bang or proportional controllers.
There are several ways to dump momentum, besides gimbaled
thrusters, which may be available on spacecraft 20, such as
magnetic torquers 46, chemical thrusters, solar panels 28 and
momentum or reaction wheels 45. Momentum dumping control 66 sends
signals to one or more of the momentum accumulators and dumpers so
that the appropriate amount of momentum is dumped.
The dumping of momentum by reaction wheels 45 is well known and
understood in the art.
In the case of a magnetic torquer, it can have several forms
including either a magnetic torquer coil or magnetic torquer bars.
A magnetic torquer coil would generally be mounted on the
anti-nadir face of the spacecraft in the roll-pitch plane. A
voltage would be applied to this large coil such that it produces a
magnetic dipole which interacts with the Earth's magnetic field
(which is generally directed towards the South). This combination
produces body torques which are generally only directed along the
roll axis of the spacecraft. Magnetic torquer bars would be similar
units which can produce dipoles in any axis, however for this
orbit-normal spacecraft attitude, they would also produce torques
only about the roll axis. They must be applied at the right time of
day when the roll axis is properly located with respect to the
disturbance torque which is to be counteracted (the sun is the
dominant source for this disturbance torque which rotates once per
day in the roll-yaw plane).
Regarding chemical thrusters, they traditionally can be controlled
about any axis, but also may result in higher pointing transients
(and is more expensive of a system in hardware and fuel costs).
This is traditionally used in thresholding scheme whereby a thrust
pulse would be used to dump momentum when that axis exceeds a
certain threshold. Pitch disturbances are traditionally dumped with
chemical propulsion.
Solar tacking of a solar panel can be used to dump momentum by
commanding offset angles to the solar wing drives to also produce
torques in the roll-yaw plane. Solar tacking and magnetic torquing
might only be used in conjunction with a xenon propulsion system if
the disturbance torques were exceptionally high (such as for a very
asymmetric spacecraft design).
The difficulty of integrating these multiple momentum dumping
schemes together is in accounting for their different periods of
operations and their different capabilities in different axes.
Momentum dumping control 66 can take into account the different
momentum dumping capabilities of the above-mentioned momentum
dumpers and formulate one or more methods of coordinating the
dumping of momentum off the spacecraft 20.
One method would be for the momentum dumping control 66 to predict
the amount of H.sub.z (byproduct) momentum which will be produced
by the electronic thrusters when dumping H.sub.y. During the
worst-case time of year it is the H.sub.z momentum command which
will risk saturating the capability of the electronic thrusters.
The portion of H.sub.z which exceeds the electronic thrusters'
capability can then be commanded to a solar tacking (or magnetic
torquing) controller (not shown).
Another role of momentum dumping control 66 would be to prevent
allowing the multiple dumping mechanisms to confuse or even fight
against each other. One example is shown in FIG. 10 where it is
shown that the thruster momentum produced in the xz plane is not
directed purely against the targeted disturbance torque, but is
composed of multiple skewed momentum commands. Each thruster burn
stores undesirable momentum in the momentum accumulators that will
be removed at the next burn. If both thrust burns and solar tacking
were allowed to sample the momentum accumulators to determine the
amount of momentum to dump, then solar tacking would not know how
much of the current momentum is temporary (and will be removed by a
subsequent thruster burn). The momentum dumping control 66 will be
responsible for sampling the momentum wheels and computing dumping
commands for all mechanisms.
Various techniques may be used to improve performance for the
above-described momentum dumping/stationkeeping technique. For
example, the value of .delta..sub.yz may be varied throughout the
year to minimize transient momentum storage and/or reduce required
gimbal angle .rho.. This parameter can either be controlled by the
ground (if it varies slowly through the year) via the equation:
or, with great difficulty, it can be computed autonomously on the
satellite 20, such that pitch dumping is performed possibly at one
burn only.
Another avenue for improving the momentum dumping technique is have
the pitch distribution computation more accurately predict the
roll/yaw byproduct based on burn distribution and or
.delta..sub.yz.
A third improvement would be to increase dumping capability by
increasing burn duration symmetry using a stationkeeping strategy
which controls eccentricity less tightly. Such an approach may
still be amenable to spacecraft collocation.
A fourth improvement is to adopt an alternative algorithm that may
be used such that roll transient momentum storage (between each
west and east burn) is minimized. This can reduce the momentum
storage requirement (possibly require a smaller reaction wheel) at
the cost of an increase in thruster platform gimbal range. Such an
algorithm may require changing only the distribution factors
k.sub.w and k.sub.e, which may have values ranging from 1/3to
2/3.
As seen in FIG. 9, both dumping techniques--minimized gimbal angles
and minimized transient roll--are compared. This figure regards the
scenario of dumping 5,5,10 Nms daily in roll, pitch, and yaw
respectively. The first algorithm is optimized for minimum gimbal
angle, resulting in four momentum vectors which are proportional to
their burn durations. The cost of this optimization is a 100%
increase in the roll transient momentum for the geometry shown
above (between the D.sub.1 and D.sub.2 burns for example) over the
alternative algorithm. The second algorithm is optimized for
minimum roll transient momentum as is apparent from the more
centered excursions of the roll/yaw momentum vector from the ideal
dump trajectory. The cost of this optimization is an increase in
required .rho. angle for the geometry shown above (50% increase for
the D.sub.2 burn and 100% increase for the D.sub.3 burn). The
transient momentum is increased by 5% in roll and decreased by 12%
in yaw for this example (unfortunately, the roll axis has the least
momentum margin).
Shown in FIG. 10 is another representation of the same dump
strategy which better represents the temporal location of the
torques (disturbance momentum and pitch dump torques accumulate
equally between the two nodes), such that the transient momentum is
more readily apparent.
Should one of the thrusters fail or a xenon power conditioner fail,
at least 40% more fuel will be required for stationkeeping and
momentum management. In particular, a failure of a thruster or a
xenon power conditioner requires only one burn to be performed at
each of the ascending and descending nodes along with third and
fourth burns from the thrusters at a third point in the orbit, as
described in U.S. Pat. No. 5,443,231. While a failure of a thruster
or a power conditioner requires a variation in the geometry of the
above-described stationkeeping/momentum management, the
calculations will be similar to those in the non-failure scenario
with a more generalized geometric derivation.
In summary, the present invention regards a control system for
controlling the orbital position of a spacecraft. The control
system of the present invention simultaneously controls the orbital
motion and the momentum dumping of a spacecraft which leads to
improved fuel efficiency on the spacecraft which allows a
spacecraft to perform stationkeeping solely via ion propulsion
thrusters and, thus, reduces the need for bipropellant fuel or
thrusters.
The foregoing description is provided to illustrate the invention,
and is not to be construed as a limitation. Numerous additions,
substitutions and other changes can be made to the invention
without departing from its scope as set forth in the appended
claims.
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